1. Field of the Invention
The present invention relates generally to a turbine rotor blade, and more specifically to a turbine rotor blade with a squealer tip.
Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially an industrial gas turbine engine, the turbine section includes a plurality of stages of turbine rotor blades with blade tips that from a gap with an outer shroud of the engine in which the hot gas flow passing through the turbine can leak past the blade tips. The blade tip gap leakage not only reduces the efficiency of the turbine by not impacting all of the gas flow onto the turbine rotor blades, but can cause thermal damage to the blade tips and result in shortened life for the blades.
In a high temperature turbine blade tip section, the heat load is a function of the blade tip leakage flow. A high leakage flow will induce a high heat load onto the blade tip section. High heat loads on the blade tip can cause erosion or other thermal damage to the tip that will decrease part life or decrease engine performance. Thus, blade tip section sealing and cooling must be addressed as a single problem. In the prior art, a turbine blade tip includes a squealer tip rail that extends around the perimeter of the airfoil flush with the airfoil wall and forms an inner squealer pocket. The main purpose of using a squealer tip in a blade design is to reduce the blade tip leakage and also to provide the rubbing capability for the blade.
In the prior art, blade tip cooling is accomplished by drilling holes into the upper extremes of the serpentine coolant passages from both the pressure and suction surfaces near the blade tip edge and the top surface of the squealer cavity. In general, film cooling holes are located along the airfoil pressure side and suction side tip sections and from the leading edge to the trailing edge to provide edge cooling for the blade squealer tip. In addition, convective cooling holes are also located along the tip rail at the inner portion of the squealer pocket to provide for additional cooling for the squealer tip rail. Since the blade tip region is subject to severe secondary flow field, a large quantity of film cooling holes and cooling flow is required in order for adequate cooling of the blade tip periphery.
FIG. 1 shows a prior art rotor blade squealer tip cooling design with the secondary hot gas flow migration around the blade tip section. The squealer tip pocket is formed by the pressure side and the suction side walls and the pocket floor. Film cooling holes are shown on the pressure side wall just beneath the squealer tip edge. Cooling holes are shown on the pocket floor to discharge cooling air from the internal cooling air passage and into the squealer pocket. The airflow over the blade tip flows in a vortex pattern as indicated by the arrows. FIGS. 2 and 3 shows the pressure side film cooling hole arrangement and shape of each film cooling hole opening.
The blade squealer tip rail is subject to heating from three exposed sides which are heat load form the airfoil hot gas side surface of the tip rail, heat load from the top portion of the tip rail, and heat load from the back side of the tip rail. Cooling of the squealer tip rail by means of discharge row of film cooling holes along the blade pressure side and suction side peripheral and conduction through the base region of the squealer tip becomes insufficient. This is primarily due to the combination of squealer pocket geometry and the interaction of hot gas secondary flow mixing. The effectiveness induced by the pressure film cooling and the tip section convective cooling holes becomes very limited. Also, a thermal barrier coating (TBC) is normally used in the industrial gas turbine airfoil for the reduction of blade metal temperature. However, applying the TBC around the blade tip rail without effective backside convection cooling may not reduce the blade tip rail metal temperature. FIG. 4 shows the current prior art blade tip section cooling design with a TBC applied on the outside and the inner surface of the squealer pocket. The blade tip includes a pressure side wall 11 and a suction side wall 12, a squealer tip rail 12 on both sides that forms the pocket 13, an internal cooling air supply passage 14, a TBC 15 applied to the pressure and suction side walls and to the pocket 13, a pressure side film cooling hole 16, a suction side film cooling hole 17, a pressure side cooling hole 18 in the pocket and a suction side cooling hole 19 in the pocket. Cooling air from the internal blade cooling circuit is discharged out from the four cooling holes to provide film cooling for the walls and to cool the squealer pocket.
Several prior art references disclose turbine blades with squealer tips having cooling passages to reduce the leakage and thermal effects from the hot gas flow leakage. These include U.S. Pat. No. 5,660,523 issued to Lee on Aug. 26, 1997 and entitled TURBINE BLADE SQUEALER TIP PERIPHERAL END WALL WITH COOLING PASSAGE ARRANGEMENT in which a cooling passage arrangement in the end walls surrounding the pocket. U.S. Pat. No. 4,142,824 issued to Andersen on Mar. 6, 1979 and entitled TIP COOLING FOR TURBINE BLADES discloses straight cooling passages located in the tip wall on the suction side of the blade. U.S. Pat. No. 4,487,550 issued to Horvath et al on Dec. 11, 1984 and entitled COOLED TURBINE BLADE TIP CLOSURE discloses cooling passages in both the pressure and suction side tip walls in which both are supplied with cooling air from a common inlet passage connected to the inner blade cooling passage circuit.
All of the above cited references disclose cooling passages formed within the blade tip wall to provide cooling for the wall and to discharge cooling air into the tip gas. However, these references do not change the momentum of the cooling air flowing through the cooling channels to increase the heat transfer rate coefficient, or inject the cooling air in a certain direction to limit mixing of the cooling air with the hot gas flow across the gap in order to form a well defined film sub-boundary layer on the external surface for the reduction of the external heat load onto the blade pressure and suction tip rail as does the blade tip cooling passages of the present invention.
It is therefore an object of the present invention to provide for a turbine blade tip with a cooling passage arrangement that will reduce the metal temperature of the blade tip in order to increase the part life.
It is another object of the present invention to provide for a turbine blade tip with a cooling passage arrangement that will reduce the leakage flow across the tip gap in order to increase the turbine efficiency.
It is another object of the present invention to provide for a turbine blade tip with a cooling passage that will inject the cooling air onto the tip rail wall at a smaller angle than would the prior art straight cooling holes.